The present invention relates to the general field of gas turbine engines, and more particularly to the field of high-pressure turbine nozzle-vane bands for a gas turbine engine.
A gas turbine engine typically includes a nacelle which forms an opening for admitting a determined flow of air towards the engine proper. Generally, the engine includes a compression section for compressing the air admitted into the engine, and a combustion chamber in which the air compressed in this way is mixed with fuel and then burnt. The gases generated by said combustion are then directed towards a high-pressure turbine before being exhausted.
The high-pressure turbine conventionally includes one or more rows of turbine vanes spaced apart circumferentially all around the rotor of the turbine. It also includes a nozzle assembly enabling the flow of gases from the combustion chamber to be directed towards the turbine vanes at an appropriate angle and speed so as to rotate the vanes and the rotor of the turbine.
The nozzle assembly generally comprises a plurality of guide vanes which extend radially between bottom and top annular bands and which are spaced circumferentially relative to one another. The vane bands thus come directly into contact with the hot gases from the combustion chamber. They are subjected to very high temperatures and therefore need to be cooled. The ever increasing temperatures at the outlets of combustion chambers, and the use of chambers having two heads so as to further increase the performance of engines are leading to higher and higher temperatures in the vicinity of the bands. The increasing temperature stresses at the vane bands mean that the techniques used to cool them must be reconsidered.
A cooling device for gas-turbine nozzle bands is known from American patent U.S. Pat. No. 5,197,852. The device comprises, in particular, an internal circuit provided inside the band to enable a cooling fluid to flow through the band and cool said band. In addition to the internal circuit, a thermal-barrier-forming-coating is placed on the side of the band bordering the gas stream, and extends from a zone situated between the vanes as far as the downstream end of the band so as to reduce the temperature gradient between the two sides of the band.
The cooling device of the nozzle band described in that document can turn out to be insufficient, in particular downstream from the guide vanes in the slipstream of their trailing edges where burns can appear. In addition, since the thermal barrier provided is deposited on the throat surfaces of the vanes, it can affect the throat section of the nozzle and degrade the performance of the high-pressure turbine. The zone to be covered by the thermal-barrier-forming coating is also difficult to access (in particular in the channels between vanes), thus leading to an increase in the cost of making the band.
The present invention thus seeks to mitigate such drawbacks by proposing a nozzle-vane band including a cooling device to protect the band thermally in a region in which other cooling techniques cannot be used. It also seeks to provide a nozzle band having a cooling device that does not disrupt the throat section of the guide vanes and that does not require a cooling circuit that is inside the band. It also seeks to provide a nozzle band having a cooling system that is not particularly difficult to install. Finally, it seeks to provide a high-pressure turbine nozzle including at least one band of the invention.
To this end, the invention provides a high-pressure turbine nozzle-vane band for a gas turbine engine, the band comprising an inside surface supporting at least one guide vane having a trailing edge that is directed towards a downstream end of the band, and an outside surface, opposite the inside surface, from which a flange extends radially, defining firstly, upstream from the flange, a passage for cooling-air, and secondly, downstream from the flange, a cavity, wherein the inside surface of the band is provided, between the trailing edge of the guide vane and the downstream end of the band, with a coating forming a thermal barrier enabling a temperature gradient generated in the band by the air spinning in said cavity to be increased.
In this way, the presence of the thermal-barrier-forming coating enables the band to be protected from burns which may appear downstream from the guide vanes, in the slipstream of their trailing edges.
So as not to degrade the aerodynamic performance of the high-pressure turbine, the thermal-barrier-forming coating has a surface which is substantially flush with the inside surface of the band upstream from the thermal barrier.
The outside surface of the band advantageously includes spoiler projections extending between the flange and the downstream end of the band so as to increase the temperature gradient generated in the band and thus improve the effectiveness of the thermal barrier.
The spoiler projections can be in the form of ribs that are substantially parallel or inclined relative to the axis of the turbine, or in the form of curvilinear ribs or even studs.